Most relatively small missiles in use today are propelled by solid fuel rockets as opposed to, for example, turbojet engines. The selection of a solid fuel rocket as a propulsion device has been largely dictated by two factors. First, in many instances, a turbine engine cannot be fabricated sufficiently economically as to compete with a solid fuel rocket engine. Secondly, in small size missiles, i.e., those having a relatively small diameter measured on the order of about six inches or less, it has heretofore been extremely difficult to manufacture an efficient turbojet engine that will fit within the envelope required for the propulsion unit for such a missile.
As a consequence of the resulting use of solid fuel rocket engines, some degree of control of the missile flight path or trajectory is lost over that which is available were it possible to propel the missile by a gas turbine engine whose output can be readily varied. Further, even if the gas turbine engine operates relatively inefficiently, the use of such an engine greatly extends the range of the missile.
Recently, in order to overcome the difficulties attendant the use of rocket engines, the assignee of the instant application has produced a gas turbine engine having a diameter of about six inches. These engines are disclosed in U.S. Pat. No. 4,794,754 issued Jan. 3, 1989 to Shekleton, et al and have been extremely successful in meeting the challenge of providing a propulsion unit for small diameter missiles. However, because the diameter of such an engine in effect sets the minimum frontal area for the missile, such engines have only been capable of use with missiles having diameters of six inches or more. Thus, while such engines have fulfilled a substantial need, there remains a need for an even smaller diameter, gas turbine, thrust producing engine; and the present invention is directed to fulfilling that need.